The present invention relates generally to gas turbine engine blade platforms and, more particularly, to aircraft gas turbine engine fan blade platforms.
An aircraft turbofan gas turbine engine includes a fan assembly having a plurality of circumferentially spaced apart fan blades extending radially outwardly from a rotor disk. The fan assembly typically includes a plurality of circumferentially spaced apart fan blades each having a dovetail root disposed in a complementary, axially extending dovetail groove or slot in a perimeter or rim of a rotor disk. A spinner is mounted to a front end of the fan assembly to provide smooth airflow into the fan. A radially inner flowpath boundary for the airflow channeled between the blades is provided typically by platforms at the blade roots which circumferentially abut each other between adjacent fan blades. Blade platforms may be used between compressor and turbine blades.
One current platform design uses a resin transfer molding (RTM) method design which is costly to manufacture and require several pre-forms to be assembled in a mold then injected with resin. There is also an extended cure time for the part. Metallic and composite fan blade platforms have been made and designed. There is a need for lightweight, more easily manufactured, and less expensive fan blade platforms.